The present disclosure relates generally to turbine engines and, more specifically, to a combustion liner of a combustor assembly that has an improved film cooling hole arrangement.
In a gas turbine engine, air pressurized in a compressor is mixed with fuel in a combustor to generate hot combustion gases. Energy is initially extracted from the gases in a high pressure turbine (HPT) that powers the compressor, and subsequently in a low pressure turbine (LPT) that powers a fan in a turbofan aircraft engine application, or powers an external shaft for marine and/or industrial applications. Generally, engine efficiency increases as the temperature of combustion gases is increased. However, the increased gas temperature increases the operating temperature of various components along the gas flowpath, which in turn increases the need for cooling such components to facilitate extending their useful life.
For example, known combustors include a combustion liner that requires cooling during operation of the gas turbine engine. Known turbine nozzles include hollow vanes which also require cooling. In at least some gas turbine engines, flowpath components exposed to hot combustion gases are cooled using compressor bleed air. For example, at least some known components channel the compressor bleed air through film cooling holes defined within the combustion liner or nozzles. In the combustion liner specifically, the film cooling holes are typically arranged in rows that extend transversely relative to a main flow axis of the gas turbine engine, and film cooling holes in adjacent rows are offset from each other in a staggered configuration. However, the natural swirl of the hot combustion gases channeled through the combustor can result in hot streaks forming along the combustion liner, thereby reducing the service life of the combustor.